1. Field of the Invention
The present invention relates generally to spacecraft deployment systems and, more particularly, to a deployment sequencer for controlling the deployment of a secondary structure.
2. Description of the Related Art
When a satellite is launched on a spacecraft, the external equipment of the satellite is stowed in a retracted position to conserve space and to avoid damage to the equipment. Once the satellite is in orbit, stowed equipment is deployed from the satellite to an operative position. Some of the equipment is designed to deploy in multiple stages to facilitate more compact stowage on the spacecraft. One example of a piece of equipment with multiple deployment stages is a solar panel array including reflector panels.
In a solar panel array, the solar panels are pivotally coupled end-to-end and folded accordion style for stowage. Each of the solar panels has a pair of reflectors, each pivotally coupled to an edge that is not coupled to an adjacent solar panel. To store the solar panel array, the reflector panels are folded over their associated solar panel, and the solar panels are folded accordion style into a panel stack. Once the satellite is in orbit, the components are unfolded in reverse order to deploy the solar panel array.
During the primary deployment of the solar panel array, the solar panels are unfolded away from the spacecraft until the panels are in substantially planar alignment. At this point, the solar panels have deployed far enough that the reflector panels coupled to one of the solar panels will not collide with the reflector panels coupled to the adjacent solar panel as the reflector panels are rotated into position during the secondary deployment. The secondary deployment stage is triggered by the release of the reflector panels from the stowed position and is completed when the reflector panels have rotated into their deployed position.
A basic two-stage mechanical deployment system, such as that described above, uses a trigger placed near the end of the path of travel of the equipment during primary deployment to operate a lever system that unlatches and starts the secondary deployment. The trigger is spring loaded so as to require the application of an activation force by a main deployment spring in order to cause the lever system to operate. Due to variations in tolerances and assembly, the activation force can vary significantly. The trigger cannot be so sensitive that it can accidentally cause the lever system to operate. Conversely, the trigger cannot be so stiff that the main deployment spring has difficulty applying a sufficient activation force. The main deployment spring must therefore be sized such that it can provide an activation force sufficient to overcome the spring loading of the trigger while providing a margin for safety. However, an oversized deployment spring stores excessive energy that is difficult to control and causes excessive strain on other parts of the system. Further, the placement of the trigger near the end of the path of travel of the equipment during primary deployment results in the activation force of the main deployment spring being at a minimum.
One alternative method of triggering the deployment of the reflector panels is to use an electronic, pyrotechnic or thermal actuator. Once the solar panels are deployed, a signal is transmitted from the satellite to the actuator to trigger the release of the reflector panels. Although these types of actuators can be effective for triggering the release of the reflector panels, the actuators add weight to the solar panel array, thereby increasing the size of the spring necessary to deploy the solar panel. In addition, this alternative requires additional slip-rings on the solar wing drive to transmit power across the joint to the actuator. Moreover, the actuators and additional slip-rings introduce additional failure modes into the system which ultimately reduce the system reliability.
Another alternative for deploying the reflectors is to link the corners of the reflector panels together with lanyards. In this alternative, the lanyards interlock the reflector panels in a semi-closed position until the solar panels are almost fully deployed, and the reflector hinge lines are linearly aligned. Although this alternative is simple and lightweight, the lanyards allow some relative movement of the reflector panels and, therefore, do no eliminate collisions between the reflectors panels.
In yet another alternative triggering method, a trigger is used near the deployment stop of the main hinge. When the main hinge has deployed and is near the deployment stop, the trigger is activated and a system of springs and levers release a latch which deploys the reflector panels. As with the use of actuators, this method adds complexity to the system, increase the weight of the satellite and the solar panel array, and reduces the reliability of the deployment system. Additionally, the pull of the trigger takes away from the torque available from the main spring at the end of deployment when the torque is at a minimum.
Therefore, there is a need for an improved deployment sequencer for triggering a secondary deployment stage without significantly increasing the weight of the satellite and the deployed equipment, significantly increasing the power required to complete deployment, or significantly reducing the reliability of the deployment system.